Single skin combustor heat transfer augmenters

ABSTRACT

A combustor for a gas turbine engine comprises a single skin liner defining a combustion chamber. The single skin liner has an inner surface facing the combustion chamber and an outer surface exposed to a coolant flow discharged in a plenum extending from the outer surface of the single skin liner to the engine casing. Cooling holes extend through the single skin liner. Open flow guiding channels are provided on the outer surface of the single skin liner, the open flow guiding channels being uncovered and aligned with the flow of air over the outer surface of the single skin liner.

TECHNICAL FIELD

The application relates generally to gas turbine engines and, more particularly, to single skin combustor liner cooling.

BACKGROUND OF THE ART

Compared to double or multi-skinned combustors, a single skin design has the potential to be lighter in weight and hence lower in cost. However, current effusion cooled liner designs are limited in efficiency due to manufacturing constraints such as hole size and angle. Therefore, without increasing cooling air consumption, additional heat removal is a challenge. In aviation gas turbine engines, it is desirable that the amount of air supplied for cooling combustor walls be minimized in order not to negatively affect the overall performances of the engine. This poses challenges to meeting the durability requirements of single skin combustor walls, because the reduction in combustion wall cooling air may lead to unwanted material oxidation, thermal mechanical fatigue and/or thermal wall buckling due to thermal gradients. Particularly in small aero gas turbine engines, the total amount of air available for combustor wall cooling within the gas turbine thermodynamic cycle can be limited, especially where rich-burn combustion is sought. Therefore, it is a challenge to optimize the combustor wall cooling while still meeting the durability requirements of single skin combustors.

SUMMARY

In one aspect, there is provided a single skin combustor for a gas turbine engine, the single skin combustor comprising: a single skin liner defining a combustion chamber, the single skin liner having an inner surface exposed to the combustion chamber and an outer surface exposed to a flow of air, the outer surface of the single skin liner being an outermost surface of the combustor, cooling holes extending through the single skin liner, and open flow guiding channels provided on the outer surface of the single skin liner, the open flow guiding channels being uncovered and aligned with the flow of air over the outer surface of the single skin liner.

In another aspect, there is provided a gas turbine engine comprising a gas generator case, a combustor disposed within the gas generator case, the combustor having a single skin liner circumscribing a combustion chamber, the single skin liner and the gas generator case defining therebetween a plenum, the single skin liner having an outer surface exposed to a flow of air discharged in the plenum and an inner surface exposed to combustion gases in the combustion chamber, cooling holes defined in the single skin liner, the cooling holes fluidly liking the plenum to the combustion chamber, and heat transfer augmenters provided on the outer surface of the single skin liner, the heat transfer augmenters including ribs configured to redirect the flow of air from a first direction to a second direction.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-section of a gas turbine engine;

FIG. 2 is an enlarged cross-section view of a portion of the combustor of the engine shown in FIG. 1 and illustrating a single skin liner with a combination of different heat transfer augmenters for guiding the incoming flow of cooling air and locally increasing heat exchange surfaces on the cold outer side of the liner;

FIG. 3 is an enlarged cold side view illustrating the heat transfer augmenters projecting from the outer surface of the single skin combustor liner; and

FIG. 4 is a schematic conceptual isometric view illustrating a flow field as modulated by a set of fins/ridges provided immediately downstream from a jet impact site on the cold outer surface of the single skin combustor liner; the impinging jet nozzle is fictitious and provided to illustrate the concept not an actual engine design .

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 10 of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan 12 through which ambient air is propelled, a compressor 14 for pressurizing the air, a combustor 16 in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section 18 for extracting energy from the combustion gases.

The combustor 16 is a single skin combustor. That is the combustor 16 has a single skin liner. According to one embodiment, the single skin liner comprises a radially inner liner 20 a and a radially outer liner 20 b concentrically disposed relative to a central axis of the engine and defining therebetween an annular combustion chamber 22. The radially inner and radially outer liners 20 a, 20 b may each be made from a single sheet of metal with through holes defined therein for cooling purposes. In contrast, double or multi-sheet liners have gaps of cooling air made by sandwiching two or more sheets of metal or mounting heat shields on the inner surface of a liner to maintain some form of air gap through which cooling air may be guided to cool the innermost skin of the liner.

A plurality of circumferentially spaced-apart nozzles (only two being shown at 28) are provided at the dome end of the combustor 16 to inject a fuel/air mixture into the combustion chamber 22. Igniters (not shown) are provided along the upstream end portion of the combustion chamber 22 downstream of the tip of the nozzles 28 in order to initiate combustion of the fuel/air mixture delivered into the combustion chamber 22. The inner and outer liners 20 a, 20 b define a primary zone of the combustion chamber 22 at the upstream end thereof, where the fuel/air mixture provided by the fuel nozzles is ignited. The primary zone is generally understood as the region in which the fuel is burned and has the highest flame temperature within the combustor 16. The combustor 16 also has a secondary zone, which is the region characterized by first additional air jets to quench the hot product generated by the primary zone; and a dilution zone corresponding to the region where second additional jets quench the hot product and profile the hot product prior to discharge to the turbine section 18.

The combustor 16 is mounted in a plenum 17 circumscribed by an engine casing 26 (e.g. a gas generator case). The plenum 17 extends from the single skin liner of the combustor 16 to the engine casing 26. In other words, the single skin liner is an outermost surface of the combustor 16. The single skin liner is free of coverage in the plenum 17 (it is not surrounded/covered by any flow guiding sleeve to form an air gap like in a double skin design). The plenum 17 is supplied with compressor bleed air from the compressor 14. Compressor exit tubes such as the one shown at 21 in FIG. 4, can be used to direct high momentum air cooling jets onto the outer surface 36 of the combustor liner.

As illustrated in FIG. 2 in relation with the inner single skin liner 20 a, a plurality of cooling holes 30 are defined in the inner and outer single skin liners 20 a, 20 b for allowing air in the plenum 17 to flow through the liners 20 a, 20 b, thereby picking up heat therefrom, and to then form a protective film of cooling air over the inner or combustion facing surface 32 of the liners 20 a, 20 b. The amount of cooling that is required is typically higher in the primary zone than in the secondary zone in the combustor. Dilution holes 33 also extend through the inner and outer liners 20 a, 20 b in the secondary zone of the combustor 16. The dilution holes 33 are not to be confused with the cooling holes 30. The dilution holes 33 are used to introduce dilution air into the combustion zone of the combustor 16. The dilution air quenches the flames so as to control the gas temperature to which the turbine hardware downstream of the combustor will be exposed. The quenching also reduces the level of NOx emissions in the engine exhaust. The dilution holes 33 are generally far smaller in number than the cooling holes 30, and each dilution hole 33 has a cross-sectional area that is substantially greater than the cross-sectional area of one of the cooling holes 30. The dilution holes 26 are typically arranged in a circumferentially extending row.

The compressor bleed flow discharged in the plenum 17 typically produces an uneven profile which causes some regions to exhibit flow stagnation or recirculation and others to experience flow separation. Specifically, for combustor cold side cooling, the heat removal occurs mainly at the surface by means of forced convection. Stagnation zones have a detrimental effect on this form of heat transfer. Such detrimental effects can be minimized and cooling efficiency thus improved by augmenting cold side surface flow characteristics and enhancing heat transfer. This can be accomplished by providing positive and/or negative material features on the outer or cold side surface 36 of the single skin liners 20 a, 20 b to channel and guide cooling flows into hotter regions for enhanced cooling. For instance, heat transfer augmenters, including various types of flow guiding structures and flow heat transfer enhancement features, may be strategically positioned on the cold outer surface or back side 36 of the inner and outer liners 20 a, 20 b to guide the coolant flow in open channels 37 (FIG. 4) to most thermally solicited regions of the inner and outer liners 20 a, 20 b and provide additional cooling thereat by increasing the surface area available for convection cooling. As will be seen hereinafter, the open channel concept of cold side heat transfer augmenters allow incident flow to approach the combustor as it normally does but once it nears the outer surface 36, diverts it towards much needed areas. Once it reaches the desired areas, heat transfer promoting features such as turbulators (e.g. pin fins) can be used to create a larger heat transfer benefit. Open-channel flow may be defined as a type of flow where the fluid is partially contained by a hard surface but also has a free surface exposed (example: a river). Closed-channel flow is where the fluid does not have any free surface and is fully contained by the solid walls (example: a pipe). In this case, the heat transfer augmenters are formed by a series of small ribs that channel surface flow similar to a moat or a river. Since the surfaces features (or channels) are so small compared to the casing around the combustor, it is similar to open-channel flow. If one were to take a cross-sectional tracing of the wall boundaries containing the flow, 90 degrees to the streamline, an open-channel would have an open profile (a line) whereas a closed-channel would have a closed profile (a circuit).

Referring more particularly to the embodiment shown in FIGS. 2 and 3, it can be seen that the heat transfer augmenters may comprise an upstream row of circumferentially distributed ribs 34 oriented to guide the upcoming flow of cooling air to an intermediate band of circumferentially distributed pin fins 39 positioned upstream of a downstream circumferential row of ridges or fins 38, which is in turn disposed upstream of the dilution holes 33. It is understood that the illustrated combination of heat transfer augmenters constitute only one possible combination and arrangement of back side heat transfer augmenters that could be used to guide the coolant flow and enhance heat transfer on the cold outer side surface 36 of the single skin inner and outer liners 20 a, 20 b. In the illustrated embodiment, the ribs 34, the pin fins 39 and the fins 38 are all disposed in the primary zone of the combustor 16. However, it is understood that similar cold side heat transfer augmenters could be provided in the secondary zone of the combustor as well.

The ribs 34 are aligned with the flow of air on the outer surface 36 of the combustion liners 20 a, 20 b. The ribs may be curved so that as the flow of air changes direction, the ribs remain aligned. As best shown in FIG. 4, the adjacent ribs 34 form open flow guiding channels 37 on the outer surface of the liners to guide the flow to a desired location thereon. Depending on the flow direction of the incoming air and on the location of the hotter regions requiring additional cooling, the ribs 34 may be angled to the axial direction or curved. Also, adjacent ribs 34 may be non-uniformly spaced to vary the amount of cooling air being channeled to differently thermally solicited regions of the liner.

The pin fins 39 are typically provided on the hotter regions of the liners 20 a, 20 b where additional cooling is required. The pin fins 39 are disposed to receive the flow of air guided by the upstream set of ribs 34. Other turbulators, such as trip strips, could be used in place or in combination with the pin fins 39. The downstream fins 38 can performed a flow guiding function as well as heat exchange promoting function. As the ribs 34 they can be curved or angled to direct the flow of cooling air coming from the pin fins 39 in any desired direction over the outer surface 36 of the liners 20 a, 20 b.

The ribs 34, the pin fins 39 and the fins 38 may each be provided in the form of free-standing protrusions integrally projecting from the outer surface 36 of the radially inner and outer single skin liners 20 a, 20 b. Each of these features may be integrally formed on the outer surface 36 of the liner by means of additive manufacturing or other suitable manufacturing processes. According to one embodiment, the cold side ribs 34, the pin fins 39 and the fins 38 are obtained as an extension of a base metal of the single skin liner by laying down successive layers of the base metal onto the outer surface of a sheet metal substrate.

It is understood that the ribs 34, the pin fins 39 and the fins 38 could each have various configurations and geometries. For instance, as shown in FIG. 4, the ribs 34 could include a combination of straight and curved ribs 34 a, 34 b to provide for a better distribution of the cooling air over the outer surface 36 of the single skin liners 20 a, 20 b. Such ribs could be positioned to receive and guide the flow of cooling air discharged by compressor exit tubes 21 onto the outer surface 36 of the single skin liners 20 a, 20 b. The straight ribs 34 a are centrally disposed relative to the jet impact site, whereas the curved ribs 38 b curve away from the straight ribs 34 a on opposed lateral sides thereof. As can be appreciated from FIG. 4, the flow of incoming jet is captured and the edges (lateral portions of the jet) are redirected sideways by the curved lateral ribs 34 b. The flow speed is maintained for a longer distance from the jet impact site, which increases heat transfer.

The heat transfer augmenters (ribs 34, the pin fins 39 and the fins/ridges 38) could be distributed on a partial surface of the single skin liner or over a full surface thereof. The heat transfer augmenters are distributed so as to provide for a uniform temperature distribution all around the combustor liners 20 a, 20 b. For instance, the density of pin fins 39 (or other suitable turbulators) could be greater in hot spot regions and less in cooler regions of the combustor 16. Also, a greater concentration of heat transfer augmenters can be provided in certain regions of the combustor 16 where it is desirable to limit the quantity of cooling air flowing into the combustor because the cooling air may have a detrimental effect on the overall combustion process. For instance, in some applications, it might be desirable to cut down on the amount of cooling air directed into the primary zone of the combustor 16 in order to maintain a rich fuel/air mixture ratio. This may be achieved by reducing the density of cooling holes 30 in the primary zone and correspondingly increasing the density of heat transfer augmenters in this same primary zone so as to compensate for the reduced number of cooling holes.

The heat transfer augmenters can be of uniform or non-uniform height. Also, it is understood that a combination of different shapes and kind of heat transfer augmenters could be provided on the cold outer surface 36 of a same single skin liner 20 a, 20 b. In fact, various combinations of sizes, distributions and dimensions are possible.

In use, the compressor bleed air is discharged into the plenum 17 with significant momentum. A significant portion of this momentum is converted to static pressure upon encountering the combustor liner, prior to entering the combustion chamber 22. The heat transfer augmenters (the ribs 34, the pin fins 39 and the fins 38) are passive features that utilize some of the air momentum before it gets converted. The air is captured in the open channels 37 formed between the ribs 34 and is guided and redistributed over the outer surface 36 of the liners 20 a, 20 b to promote a more uniform temperature thereover. It allows to augmenting the local flow such that the heat transfer coefficient is higher where it is needed. The air flowing through the guiding channels 37 between the ribs 34 is directed at least in part to the pin fins 39, which are provided in the hot spot regions to enable higher heat transfer efficiency in the most thermally solicited regions of the combustor liner. In this way, the risk of having over-cooled and under-cooled regions may be reduced. It provides for a more uniform temperature distribution around the combustor liner. The cooling air leaving the pin fins 39 is then received and guided by the fins 38. As the air flows in the open channels defined between adjacent fins 38, it picks up heat from the fins . As mentioned hereinbefore, the fins may be curved and/or appropriately oriented relative to the flow of cooling air to guide the air to a desired location before it reaches the row of dilution holes 33.

From the foregoing, it can be appreciated that the cold side heat augmenters allow the designer to reduce hot spots, improving both the thermal mechanical and oxidation life of the part. This is all done without the need to increase cooling air consumption; thereby minimizing the impact on combustor and overall engine efficiency.

The air guided on the outer surface 36 of the liners 20 a, 20 b has a second opportunity to cool the liners 20 a, 20 b by flowing through the cooling holes 30. Indeed, as the air flows through the cooling holes 30, it cools the liners 20 a, 20 b by in-hole heat transfer. At its exits from the cooling holes 30, the air flows over the inner or hot combustion facing surface 32 of the liners 20 a, 20 b, thereby providing for the formation of a protective cooling film thereover. Accordingly, with the addition of the heat transfer augmenters on the cold side of the liner, the air has several opportunities to cool down the liner. Multiple usage of the same cooling air provides for improved cooling efficiency. In this way, single skin combustors may be used in high temperature applications where double skin combustor designs would have typically been retained. Also, since the heat transfer augmenters are located on the cold side of the combustor liner, they are not exposed to the hot combustion gasses and are, thus, less subject to erosion over time. This provides for a more robust design.

While cold side heat transfer augmenters have been mainly described in connection with positive material features (e.g. ridge, ribs, fins and pin fins), it is understood that they could also take the form of negative material features. For instance, dimples could be formed in the outer surface 36 of the liner to act as turbulators. The open channels between the ribs could take the form of closed bottom grooves or slots machined or otherwise suitably formed in the outer surface 36 of the liners 20 a, 20 b. Surface knurling could also be used to form heat transfer augmenter patterns of various configuration (e.g. star pattern, fan pattern) in the outer surface 36 of the liners 20 a, 20 b.

The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. For instance, the same principle could be applied to a combustor can. Any modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims. 

1. A single skin combustor for a gas turbine engine, the single skin combustor comprising: a single skin liner defining a combustion chamber, the single skin liner having an inner surface exposed to the combustion chamber and an outer surface exposed to an open-channel flow of air in a plenum circumscribed by a gas generator case of the gas turbine engine, the outer surface of the single skin liner being an outermost surface of the combustor, cooling holes extending through the single skin liner, and open flow guiding channels provided on the outer surface of the single skin liner, the open flow guiding channels being uncovered and aligned with the flow of air over the outer surface of the single skin liner.
 2. The single skin combustor defined in claim 1, wherein at least some of the open flow guiding channels define a curve to in use redirect a portion of the flow of air to a predetermined hot spot region.
 3. The single skin combustor defined in claim 2, wherein heat transfer augmenters are provided in the hot spot region, the heat transfer augmenters projecting from the outer surface of the single skin liner.
 4. The single skin combustor defined in claim 3, wherein the heat transfer augmenters are selected from the group consisting of: pin fins, fins, trip strips, dimples and knurled surfaces.
 5. The single skin combustor defined in claim 1, wherein the open flow guiding channels are defined between pairs of adjacent ribs projecting from the outer surface of the single skin liner.
 6. The single skin combustor defined in claim 5, wherein at least some of the ribs are curved.
 7. The single skin combustor defined in claim 5, wherein the ribs are non-uniformly distributed on the outer surface of the single skin liner to in use direct more air towards predetermined hot spot regions of the single skin liner.
 8. The single skin combustor defined in claim 5, wherein pin fins project from the outer surface of the single skin liner downstream from the ribs relative to the flow of air over the outer surface.
 9. The single skin combustor defined in claim 8, wherein a circumferential row of fins extends from the outer surface of the single skin liner downstream from the pin fins.
 10. The single skin combustor defined in claim 9, wherein a circumferential row of dilution holes is provided downstream of the circumferential row of fins.
 11. The single skin combustor defined in claim 3, wherein the heat transfer augmenters are an extension of a base metal of the single skin liner.
 12. The single skin combustor defined in claim 11, wherein each heat transfer augmenter comprises successive layers of sequentially deposit material on a sheet metal base.
 13. The single skin combustor defined in claim 3, wherein the density of heat transfer augmenters varies over the outer surface of the single skin liner.
 14. A gas turbine engine comprising a gas generator case, a combustor disposed within the gas generator case, the combustor having a single skin liner circumscribing a combustion chamber, the single skin liner and the gas generator case defining therebetween a plenum, the single skin liner having an outer surface exposed to a flow of air discharged in the plenum and an inner surface exposed to combustion gases in the combustion chamber, cooling holes defined in the single skin liner, the cooling holes fluidly liking the plenum to the combustion chamber, and heat transfer augmenters provided on the outer surface of the single skin liner, the heat transfer augmenters including ribs configured to redirect the flow of air from a first direction to a second direction.
 15. The gas turbine engine defined in claim 14, wherein the ribs are oriented towards a predetermined hot spot region of the single skin liner.
 16. The gas turbine engine defined in claim 15, wherein the heat transfer augmenters further comprise turbulators in the predetermined hot spot region, the turbulators projecting from the outer surface of the single skin liner.
 17. The gas turbine engine defined in claim 16, wherein the turbulators are selected from the group consisting of: pin fins, trip strips, dimples and knurled surfaces.
 18. The gas turbine engine defined in claim 15, wherein the ribs are curved.
 19. The gas turbine engine defined in claim 14, wherein the heat transfer augmenters further comprise a plurality of pin fins downstream from the ribs relative to the flow of air over the outer surface of the single skin liner, and a set of fins downstream from the pin fins, the pin fins and the fins projecting into the plenum.
 20. The gas turbine engine defined in claim 14, wherein the single skin liner is provided in the form of a metal sheet, and wherein the heat transfer augmenters include free standing ribs projecting from the outer surface of the metal sheet. 